The present invention relates generally to gas turbine engines, and, more specifically, to bleed systems therein.
A turbofan aircraft engine includes a fan mounted inside a surrounding nacelle, and is driven by a low pressure turbine (LPT). An inner portion of air channeled through the fan enters a core engine in which the air is pressurized in a high pressure compressor (HPC) and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in a high pressure turbine (HPT) that drives the compressor.
The outer portion of fan air bypasses the core engine through an annular bypass duct. The pressurized air discharged from the bypass duct provides a majority of propulsion thrust of the engine for powering an aircraft in flight.
In large turbofan engines, additional power is generated by including a low pressure or booster compressor behind the fan and in front of the HPC of the core engine. The booster compressor typically includes multiple axial stages which increase pressurization of the fan air delivered to the HPC, which in turn includes multiple axial stages further increasing the pressure of the air provided to the combustor.
The typical turbofan aircraft engine is configured for operating over a flight envelope including idle, takeoff, climb, cruise, runway approach, and landing in which the power output of the engine correspondingly varies. For example, the multiple axial stages of the booster and high pressure compressors must be designed and operated for obtaining a suitable stall margin over the entire operating range. For maximum power operation of the engine, the compressors are operated at maximum airflow and maximum pressurization, with a suitable stall margin.
However, at flight idle operation during landing approach of the aircraft the engine produces relatively low power, and the HPC requires correspondingly less airflow therethrough. In order to maintain efficient operation of the engine at this part power condition, and maintain a suitable stall margin in the HPC, a portion of the pressurized booster compressor air is typically bled from the engine and dumped into the fan bypass duct.
Accordingly, a booster bleed system is typically incorporated in large turbofan aircraft engines for selectively bleeding a portion of the booster discharge air when desired for maintaining efficient operation of the engine, including suitable compressor stall margin.
The typical booster bleed system is relatively large and relatively complex and is located between the booster and high pressure compressors. For example, the turbofan engine includes a fan frame disposed between the two compressors. The frame includes a row of struts extending radially outwardly through the fan bypass duct to support the fan nacelle.
The frame also includes a center structural hub having a row of low transition ducts alternating between the inner ends of the struts for providing flow continuity between the outlet of the booster compressor and the inlet of the HPC. The hub also includes one or more bearing supports which contain bearings for supporting the fan drive shaft that joins the fan to the LPT. The rotor blades of the booster compressor are also joined to the fan drive shaft.
In a large turbofan engine, the fan frame is correspondingly large, with a correspondingly large center hub in which the typical booster bleed system may be incorporated. However, incorporation of that bleed system: correspondingly requires inlet apertures in the hub for bleeding booster air. Outlet apertures are also required in the hub for channeling the bleed air into corresponding outlets in the fan bypass duct.
Any hole or aperture placed in the structural hub of the fan frame interrupts the structural integrity thereof and correspondingly requires strengthening of the hub which typically increases size and weight of the fan frame. The bleed system also requires multiple inlet valves or doors and corresponding actuating mechanisms for selectively opening and closing the bleed doors when required during operation of the engine.
The bleed system mounted inside the typical fan frame of a large turbofan engine increases the cost of manufacture of the engine, increases weight of the engine, and correspondingly decreases overall efficiency of the engine.
In the continuing development of high-bypass turbofan aircraft gas turbine engines, reductions in size and weight of the engine, without corresponding reductions in power rating are desired. In one engine undergoing development, the fan frame includes a relatively small center hub which lacks available space for introducing a conventional booster bleed system.
Furthermore, components adjoining the fan frame have limited available space for mounting those components themselves, without the additional complication of introducing a suitable booster bleed system.
Accordingly, it is desired to provide a turbofan aircraft engine with an improved booster compressor bleed system being relatively compact and simple, and having a low profile for being integrated into available space in the engine.